A gas turbine engine generally includes a fan and a core arranged in flow communication with one another. Additionally, the core of the gas turbine engine generally includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gasses through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere.
It is desirable to increase an overall pressure ratio of the gas turbine engine in order to increase an efficiency of the gas turbine engine. The overall pressure ratio refers generally to a ratio of a pressure measured at a forward end of the compressor section to a pressure measured at an aft end of the compressor section. However, as the overall pressure ratio increases, a temperature of the compressed air also increases. The materials used to construct rotor blades and/or stator vanes in the compressor section typically are not designed to withstand heightened temperatures that can accompany an increased overall pressure ratio.
Accordingly, a device for cooling compressed air in a compressor section of a gas turbine engine would be useful.